1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
In a turbine, the rotor blades are exposed to different stress loads than the stator vanes. Because the rotor blades rotate (stator vanes do not rotate), the blades are under high stress loads due the centrifugal force of rotation. A rotor blade is thick in the lower span and tapers off in the direction toward the tip with the thinnest section being located at the tip. The upper span of the blade will thus have the lowest mass to carry while the lower span near to the platform will have the highest mass to carry. All of the blade above the platform must be carried by the lower span of the blade. The highest stress loads are then found at the lower span sections. In addition, where the blade is exposed to the very high temperatures, the metal material strength decreases. Thus, the blade shape and cooling circuitry must be designed to account for both the stress loads and the thermal stress due to normal operation in the engine. This is especially important for industrial engine blades because the life cycle must be very long.
FIG. 1 shows the external pressure profile for a prior art first stage blade. As shown in FIG. 1, the forward region of the pressure side surface experiences a high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than on the pressure side. The area within the two curves to the left of the mid-chord section is at a lower work pressure 11 while the area 12 is at a high delta working pressure. This translates into more cooling air working potential toward the trailing edge than in the leading edge.
FIG. 2 shows a blade external heat transfer coefficient for a turbine rotor blade. As shown in FIG. 2, the airfoil leading edge, the suction side immediately downstream of the leading edge, as well as the pressure side trailing edge region of the airfoil experience the higher hot gas side external heat transfer coefficient than the mid-chord section of the pressure side and downstream of the suction surfaces. Point 13 is the high heat load region for the blade leading edge, point 14 is the high heat load aft section of the P/S surface, point 15 is the low Q on the pressure side (P/S) and point 16 is a high Q on the suction side (S/S). In general, the heat load for the airfoil aft section is higher than in the forward section.
FIG. 3 shows the blade mainstream gas temperature profile. As seen in FIG. 3, the maximum gas temperature occurs at around 75% of the blade span height located at point 17. This translates into a high heat load. Since the pull stress at the blade upper span is low, it allows for the blade to run at a higher metal temperature. Below the 40% blade span height, the gas temperature drops off to a lower level that results in a lower heat load on the blade. This drop-off of the gas side temperature is good for the blade creep design, especially for the lower blade region with a high blade pulling load. Point 19 is in the upper blade span in which a lower pull stress and a higher allowable metal temperature is allowed. Point 18 is at a low gas temperature which is good for stress rupture.